Wing and blade structure using pultruded composites

ABSTRACT

Tapered layers of pre-cured composite material are integrated into a tapered, highly stressed laminate structure in order to provide improved compressive strength. The pre-cured composite material can advantageously be cured under tension as pultruded material, to further augment compressive strength. The thickness of composite layers can be tapered on their termination edges by mechanically abrading, chemical abrading, or other methods. Especially preferred embodiments include aircraft structural components such as wings, wing spars, wing skins, fuselage skins, rotor blades, propellers, and propeller blades. Preferred laminates can be constructed to have at least 6, 10, 30, 50, or 100 layers of material, and can have a maximum thickness of at least 0.15, 0.25, 0.5, 1.0, or 5.0 inches.

This application claims priority to U.S. provisional application Ser.No. 61/033,337, filed Mar. 3, 2008, the disclosure of which isincorporated herein in its entirety.

FIELD OF THE INVENTION

The field of the invention is composite structures.

BACKGROUND

Due to the nature of flight, desirable aircraft structures havetraditionally had high strength-to-weight ratios (strength efficiency)and stiffness-to-weight ratios (stiffness efficiency). In the lastseveral decades, carbon composite materials were often used in aircraftstructures to improve strength and stiffness efficiency in the airframe.Although carbon composite materials do provide weight savings overtraditional aluminum structures, carbon composite materials still sufferfrom a comparatively low compression strength. In a typicalunidirectional carbon composite laminate the compression strength isapproximately 50% of the tension strength of the material. This iscaused by small amplitude waviness in the unidirectional fibers. Thesesmall eccentricities in the fibers promote micro buckling of thelaminate under compressive loads.

Aircraft wing and blade structures in particular see high bendingstresses due to the cantilevered configuration of their structure, andthe thin sections required for aerodynamic performance. The bendingcreates high compression loads in the upper surface of a cantileveredwing or blade structure. Thus, although composite materials do increasethe structural efficiency (ratios of strength to weight or stiffness toweight) of aircraft structures over a typical aluminum structure, thereis still a large performance gap that can be bridged by increasing thecompression strength of the composite laminate.

It is known in the composite materials industry that pultrudedcomposites exhibit significantly higher compression strength thantypical fibers pre-impregnated with resin (pre-preg in the industryvernacular) in autoclave cured laminates. The pultrusion process oftensioning fibers and curing them under tension raises the compressionstrength of the material by over 60%. Pultruded composites also allowlower resin content and therefore a higher fiber volume fraction than acomparable pre-preg structure. Higher fiber volume fractions also leadto higher composite material stiffnesses and strength per unit weight.

FIG. 1 depicts a typical prior art process for making pultrusions. Thisfigure is adapted from “Composite Airframe Structures: Practical DesignInformation and Data”, by Michael C. Y. Niu, Hong Kong Conmilit PressLtd., 2005. One or more spools or other sources 102, 104, 106 of tape orother composite material comprising fibers unreel material into one ormore wet-out stations or resin tanks 110, 120. The material is thenpulled in tension by a pull station 130 before being cured into shape ata heated die station 140 powered by a power source 142. Finishedmaterial 150 leaves the die 140 with largely constant cross-section. Thefinished material 150 comprises pre-cured, pultruded composite fibers. Apre-cured composite material is pultruded when it is cured or formedunder tension.

The Niu book, as well as all other extrinsic materials discussed herein,is incorporated by reference in their entirety. Where a definition oruse of a term in an incorporated reference is inconsistent or contraryto the definition of that term provided herein, the definition of thatterm provided herein applies and the definition of that term in thereference does not apply.

The value of pultrusions is recognized in the industry; they areemployed in composite structures, such as high compression strengthareas like wings. FIG. 2 is an illustration of the prior art Genesis 2sailplane made by Group Genesis™. The aircraft 200 comprises a wing 210and a fuselage 220. FIG. 3 is a section view of the wing 210. The priorart wing section 300 comprises an airfoil 310 defining the outerboundary. Disposed within the airfoil are a foam vertical stiffener 310providing bending stiffness, a number of fiberglass laminate sheets 330,332, 334, 336, and several pultruded rectangular carbon rods 320, 322,324 that run the length of the wing. Pultrusions have also been used inthe wing structures made by large military airframers. However, in priorart known to the inventor, these pultrusions have always been usedeither as continuous strips running the length of the beam, or as verysmall rods in an under-stressed structure.

Typical distributed loads on a cantilevered structure such as a wingresult in a moment that drops off rapidly from the root to the tip ofthe structure. A tapered beam structure is often used to take fulladvantage of possible weight savings where extra structure is notneeded. Currently, any tapering of a pultruded structure simply dropsoff small pultruded sections and uses the material in an under-stresseddesign, otherwise the abrupt changes in a load bearing memberscross-section are transferred to the nearby supporting matrix in toosmall an area. The resulting stress riser fails the nearby supportingmatrix material and ultimately cause a failure of the laminate.Therefore, there is still a need to employ high compression strengthcomposite pultrusions in a highly stressed, tapered laminate.

SUMMARY OF THE INVENTION

The present invention provides systems, apparatus, and methods in whichtapered layers of pre-cured composite material are integrated into atapered, highly stressed laminate structure in order to provide improvedcompressive strength. It is contemplated that the pre-cured compositematerial could advantageously be cured under tension as pultrudedmaterial, to further augment compressive strength. The thickness ofcomposite layers could be tapered on their termination edges bymechanically abrading, chemical abrading, or other methods.

It is contemplated that the pre-cured composite material could bepultruded such that a layer in the laminate structure has a thickness ofat least 10/1000, 15/1000, or 20/1000 inch. The pre-cured compositematerial could advantageously include graphite, fiberglass, aramid, orboron fibers.

Especially preferred embodiments include aircraft structural componentssuch as wings, wing spars, wing skins, fuselage skins, rotor blades,propellers, and propeller blades. These components are often highlystressed structures and can benefit from increased compressive strength.

Preferred laminates can be constructed to have at least 6, 10, 30, 50,or 100 layers of material, and can have a maximum thickness of at least0.15, 0.25, 0.5, 1.0, or 5.0 inches. It is contemplated that one layerin a preferred laminate could have fibers predominantly aligned in onedirection, while another layer could have fibers predominantly alignedin another direction. This heterogeneous fiber layout allows fortailoring of the strength of a composite structure. In some embodiments,such as angled or kinked joints, layers in a laminate might have longaxes in different directions, at least 10°, 30°, 60°, 90°, or 120°apart.

Viewed from another aspect, the present inventive subject matter canprovide for increased compressive strength in a composite laminate byincluding a layer with pultruded fibers which vary in thickness by atleast 10%, 30%, or 60% along its length, and tapering the thickness ofthe composite laminate by at least 40%, 60%, or 80% from a maximumthickness. In more preferred embodiments, additional layers withpultruded fibers or other pultruded material could be included in thelaminate. In especially preferred embodiments, the fibers are graphitefibers, and the composite layers have a thickness of 10/1000, 15/1000,20/1000, or 25/1000 inch.

It is contemplated that laminates can have one or more layers whichadvantageously taper at a 10:1, 20:1, 40:1, or shallower slope. Further,preferred laminates can comprise additional layers of biased pre-pregtape to provide shear strength and shear stiffness.

BRIEF DESCRIPTION OF THE DRAWING

FIG. 1 depicts a typical prior art process for making pultrusions.

FIG. 2 is an illustration of the prior art Genesis 2 sailplane.

FIG. 3 is a section view of a prior art aircraft wing.

FIG. 4 depicts a preferred improved aircraft having improved wingstructure.

FIG. 5 is a cross-section of a preferred thin pultrusion strip.

FIG. 6 is an illustration of a preferred wing section structureincluding a structural box.

FIG. 7 is an illustration of the upper section structure of the wingbox.

FIG. 8 is a section cut of a laminate in the wing box upper sectionstructure.

FIG. 9 illustrates a preferred kinked joint in a laminated object.

DETAILED DESCRIPTION OF THE INVENTION

As shown in the drawing figures discussed below, pultruded fibers can beintegrated into a highly-stressed tapered beam structure such as astructurally efficient aircraft wing, rotor blade, or propeller.Pre-cured pultruded material is advantageously laminated into acomposite aircraft spar structure adding high compressive strengthmaterial to areas where it is most effective, thereby reducing aircraftweight and increasing structural capacity.

As used herein, “composite” means engineered materials comprising two ormore constituent materials. Of special relevance is carbon composites,in which carbon fiber is embedded in a matrix or resin. Alternatecomposites are also contemplated, including those containing fiberglass,ceramics, and other elements. A layer of composite material couldinclude a plurality of fibers positioned at an orientation with respectto the long axis of an object. A layer of composite material could alsoadvantageously include pultruded fibers in the form of a pultrusion. Allsuitable fibers including graphite, fiberglass, aramid, and boron arecontemplated, as well as all suitable matrix or resin materials.Similarly, all commercially viable pultrusion shapes and thicknesses arecontemplated.

FIG. 4 depicts a preferred improved aircraft comprising a fuselage 420and a wing 410 having a total span 430. The wing 410 (in solid lines) isshown in a position deflected from the undeformed wing position 412 (indashed lines). This deflection in the direction of arrow 414 is inresponse to the distributed aerodynamic and inertial loading arisingfrom flight. Further, the wing is a tapered structure, having a rootchord 432 and a tip chord 434. As used herein, a “tapered structure”means structures which have a tip chord 434 or narrow end that is 80% orless of the width of the root chord 432 or wide end. The term “slender”is used to mean structures for which the total span 430 or length is atleast four times the average chord or width.

A preferred composite wing 410 has representative fibers 416 in theupper structural member (compression cap) of the wing 410.Representative fibers 416 should be interpreted to include at least oneof graphite, fiberglass, aramid, and boron. Strain is a non-dimensionalmeasure of structural extension, expressed as the change in lengthdivided by original length. Because strain values tend to be small, theterm “micro-strain” is often used as a unit, meaning 1million (10 to the6th power) units of strain. A structure is said to be highly strainedwhen its maximum calculated ultimate strain is 1000 micro-strain ormore. For an aircraft, the maximum calculated ultimate strain in astructure such as a wing or blade usually arises with a loading from theaircraft at maximum weight multiplied by three (for an accelerationthree times gravitational acceleration) further multiplied by 1.5 (as afactor of safety).

In preferred embodiments, pultruded fibers in broad thin pultrusionstrips are incorporated into aircraft structure. A cross-section of apreferred pultrusion strip 500 is shown in FIG. 5. A plurality of fibers510, 512 are embedded in a matrix or resin 520. The pultrusion strip ispre-cured through any suitable process. A thin pultrusion strip 500 isone for which the width dimension 520 is at least six times thethickness dimension 504, the ratio of these is referred to herein as the“degree of thinness.” All suitable degrees of thinness are contemplated,including at least 6, 10, 20, 30, 50, and even 100. A broader stripallows for increased laminating productivity. Using a thin strip reducesthe step in cross-section when a pultrusion is dropped off. The minimumthickness 504 of the strip is limited by the current manufacturingcapabilities of the pultruding machine (currently about 0.02 inches),but is still three to four times thicker than a typical pre-preg tape.The strip can also be made up to 10 times the thickness of a pre-pregtape layer for increased laminate manufacturing efficiency. Thus,embodiments are contemplated which comprise a layer of pultrudedpre-cured composite material having a thickness of at least 10/1000,20/1000, 50/1000, or even 100/1000 of an inch.

It is contemplated herein that pultrusions could be advantageouslykitted in a manner similar to some pre-preg tape, comprising cutting thepultrusions to appropriate lengths and cutting for the intendedstructural planform. It is further contemplated that pultrusions mayalso be placed in the laminate with the help of laser placementmachines. As used herein, “biased pre-preg tape” refers to a tape madeof resin pre-impregnated composite fibers woven together. Individualfibers in such a weave might have angles biased relative to an axis, anexemplary angled bias being fibers of +45° and −45°.

As used herein, a “laminated object” refers to an object made withlaminates, such as a laminated composite structure. These laminatestypically comprise multiple layers or plies of composite with fibers ina resin. Individual layers or plies preferably have a plurality orfibers arranged in a predominantly similar orientation. Different layersin a laminate can have fibers at different angles. However, in somecases, a laminate may comprise only a single layer of material. An“object having a cavity” refers to either an object with boundariesdefining an internal void, including tubes, or open sections includingchannels. A single-cell or multi-cell wing or blade is an example of anobject having a cavity. A layer in the laminate has a “termination edge”if the layer does not extend to the outer boundaries of the laminate.Laminate layers are typically independent before curing and generallyflat before being formed into the desired shape, thus usually notinterwoven vertically with other layers, although an individual laminatelayer may be composed of woven fibers. Such laminate arrangements allowfor better use of the material in the flat laminate direction, and donot create bending and inter-laminar stresses due to interweaving ofplies. A laminate can have any number of layers including realistically4, 10, 50, 100, or even 500 layers. Thus, a laminate can have a maximumthickness of at least 0.05, 0.25, 0.5, 1.0, or even 5.0 inches.

As used herein, a “pultrusion based laminate” is a laminate arrangementwith pultrusions used as some or all of the layers in a laminate. Shearloads in preferred pultrusion based laminates are taken out by interlaidbiased pre-preg. It is contemplated that a plurality of pultrusionlayers can be applied between a layer or layers of biased material. Inareas where the laminate tapers, the ends of the pultrusions are sandedor otherwise abraded, thereby forming a slope to reduce the stressrisers that would otherwise occur at the abrupt end of a pultrusion. Insome laminates with interwoven biased layers there can be multiplepultruded strips with several tapers throughout the laminate. In otherstructures, the pultrusion can be up to ten times the thickness ofcomplementary biased layers used to take shear loads. The pultrusion canthen be sanded at a 10:1 slope or shallower slope such as 15:1, 25:1, or50:1 to slowly apply the stress transfer from the tapering pultrusion tothe surrounding material. The actual taper ratios will depend on thefiber and resin systems being used. Tapering can be done in both thelateral and longitudinal directions of a beam. As a beam tapers down itslength, pultrusions in a cap such as an upper compression cap can betapered as the bending moment decreases. Similarly, pultrusions can betapered across the span. In this manner pultrusions that are mosteffective in the center of the cap can be tapered off as they near theedges where primarily non-pultruded web material becomes more efficient.

Unless the context dictates the contrary, all ranges set forth hereinshould be interpreted as being inclusive of their endpoints andopen-ended ranges should be interpreted to include only commerciallypractical values. Similarly, all lists of values should be considered asinclusive of intermediate values unless the context indicates thecontrary.

In structures where the tapered beam is not straight, such as a kinkedor angled blade, the tapering pultruded strips can be interlaid withstrips in the other kink direction. Again biased material would be usedto distribute shear stresses between the pultruded layers.

FIG. 6 is an illustration of a preferred wing section structure 600 froman airplane 400. This exemplary wing section structure 600 comprises oneprimary structural box 610, advantageously constructed with a pultrusionbased laminate 620. The laminate 620 comprises a plurality of pultrudedstrips 622, 624 and biased pre-preg tape 626. In this embodiment thepultrusions are used in the upper section 612 surrounded by circle 650of the structural beam comprising the wing structural spar, while biasedpre-preg tape dominates the laminate on the side web 614 where shearforces dominate and compression material such as pultrusions is lessstructurally efficient. A detail of the upper section structure 650 isshown in FIG. 7.

The upper section structure 650 is a laminated object having layerscomprising strips of pultruded material 652, 654, 656. A strip ofpultruded material 656 running in the lengthwise direction is showntapering over its length. In this embodiment, the outside surface iskept smooth with continuous untapered pultrusions 652, 654 running alongthe outside while tapering occurs by chamfering an interior or insidepultruded layer 656. An outer ply of biased material or other weave 662covers the outer mold line surface keeping the pultrusions 652, 654, 656in place and preventing failure of the total structure due to crackingalong the pultrusions.

FIG. 8 is a section cut of a laminate 620 in the upper section structure650 showing multiple pultruded strips 624, 626, 630, 632 layered withbiased material 622, 628. A strip of pultruded material 632 is showntapering the lateral or cross ply direction.

A laminate (e.g. laminate 620) has a maximum thickness, oftencorresponding with a maximum number of laminate layers as might be foundin a wing root or blade kinked joint. In tapered structures such as anaircraft wing 410, it is contemplated that a laminate could be taperedat least 40%, 60%, or 80% from a maximum laminate thickness. Methodscontemplated for tapering the laminate include dropping off layers,tapering layers, and varying layer thickness. Independently of anytapering, it is contemplated that a layer (e.g. layer 632) in thelaminate could vary in thickness by at least 10%, 20%, or 30% along itslength. It is further contemplated that the first layer could have amaximum compressive stress before failure at any point along thelaminate is at least 70% of a maximum tensile stress before failure ofthe first layer, and the first layer varies in thickness by at least 10%along its length.

FIG. 9 illustrates a preferred kinked joint 900 in a laminated object,which should be interpreted to include a rotor blade and a propeller.The kink between two merging outer mold line edges 902, 904 comprises afirst pultruded strip 912 coming from a first direction and a secondpultruded strip 914 coming from a second direction. The first pultrudedstrip 912 is in a first layer 910 of the joint 900. It is contemplatedthat pultrusions might alternate continuity between the left and rightsides, as shown the right side pultrusions carry through on the upperlayer, and the left side carries through a lower layer 930. The firstlayer 910 has a long axis 918, while the lower layer 930 has a long axis938 in a different direction. In preferred embodiments, the anglebetween layer axes 918, 938 in a kinked joint 900 is at least 10°, 30°,60°, 90°, or 120°. The ratio of left and right continuity could beadvantageously varied depending on the relative loading of the joint. Asin the uninterrupted laminate described above, the pultrusions areinterlaid between biased plies of unidirectional tape which transfersshear loads between the pultruded layers 940. In the region of thejoint, the overall laminate thickness might increase from the additionallayers. However, some pultrusions 950 can be tapered as stressesdissipate. An outer layer of biased or woven material 960 is also shown.

Tapering of a laminate, layer, pultrusion, or laminate layer terminationedge can be accomplished in any suitable manner, including for examplemechanical, chemical, or any other type of processing. Within mechanicalprocessing are included machining, grinding, mechanically abrading, anduse of a water jet. Within the chemical category, chemical abrading isincluded, and within the category of “other” is included use of a laser.

Thus, specific embodiments and applications of using pre-cured compositematerial in a tapering, highly stressed structure have been disclosed.It should be apparent, however, to those skilled in the art that manymore modifications besides those already described are possible withoutdeparting from the inventive concepts herein. The inventive subjectmatter, therefore, is not to be restricted except in the spirit of theappended claims. Moreover, in interpreting both the specification andthe claims, all terms should be interpreted in the broadest possiblemanner consistent with the context. In particular, the terms “comprises”and “comprising” should be interpreted as referring to elements,components, or steps in a non-exclusive manner, indicating that thereferenced elements, components, or steps may be present, or utilized,or combined with other elements, components, or steps that are notexpressly referenced. Where the specification claims refers to at leastone of something selected from the group consisting of A, B, C . . . andN, the text should be interpreted as requiring only one element from thegroup, not A plus N, or B plus N, etc.

What is claimed is:
 1. A method of forming a tapered, highly stressedstructure, comprising: providing strips of pre-cured pultruded compositematerial, the strips comprising one or more fibers, and the stripshaving a width dimension and a thickness dimension, and such that thewidth dimension is at least six times the thickness dimension;constructing a laminate including at least one strip and ant least onelayer of biased pre-preg tape, the laminate having a maximum laminatethickness; and tapering a thickness of a termination edge of thelaminate such that at least one fiber in a strip is tapered.
 2. Themethod of claim 1, further comprising providing strips of pre-curedpultruded composite material such that the strip has a thickness of atleast 10/1000 inch.
 3. The method of claim 1, further comprisingproviding strips of pre-cured pultruded composite material such that thestrip has a thickness of at least 20/1000 inch.
 4. The method of claim1, wherein the pre-cured pultruded composite material comprises fibersselected from the list consisting of graphite, carbon, aramid, andboron.
 5. The method of claim 1, further comprising the step ofincluding the laminate as a structural component of an aircraft.
 6. Themethod of claim 5, further comprising the step of including the laminateas a structural component of at least one of a wing, a rotor blade, anda propeller.
 7. The method of claim 5, further comprising the step ofincluding the laminate as a structural component of a skin.
 8. Themethod of claim 1, wherein the step of constructing the laminate yieldsa maximum laminate thickness of at least 0.25″.
 9. The method of claim1, wherein the step of constructing the laminate comprises using atleast four additional strips of pre-cured pultruded composite material.10. The method of claim 1, wherein the laminate has a first strip withfirst fibers in a first direction, and the step of constructing thelaminate further comprises including a second pultruded, pre-cured striphaving second fibers in a second direction.
 11. The method of claim 1,wherein the laminate has a first strip with a first long axis and thestep of constructing the laminate comprises including a second stripwith a second long axis, and the first and second strips have long axesin different directions.
 12. The method of claim 11, wherein thedifferent directions vary by at least 10°.
 13. The method of claim 1,wherein the step of tapering comprises mechanically abrading.
 14. Themethod of claim 1, wherein the step of tapering comprises other thanmechanical or chemical abrading.